Nacelle cooling and ventilation system

ABSTRACT

A nacelle cooling and ventilation system for an aircraft gas turbine engine which includes a chin scoop and duct assembly for directing air through the engine bay where it is exhausted into the engine exhaust through an exhaust shroud eductor. The nacelle is enclosed so that effectively all of the cooling air entering the bay leaves the bay through the engine exhaust and is not dumped overboard, thereby adding to the effective thrust of the engine. An engine oil and generator oil heat exchanger is positioned in the cooling air duct into the bay so that the same air which cools the bay also cools engine and generator oil. The oil heat exchanger includes directional louvers which regulate and keep to a minimum the amount of cooling air and direct cooling air forwardly in the bay so that the entire bay area is cooled, and which can be shut as a fire control measure. The nacelle includes bay vent doors which are opened in response to a high oil temperature condition or in an engine shutoff condition on ground to provide auxiliary air flow from the bay. The doors are positionable by an actuator but can be disconnected from the actuator to pivot freely in response to a sudden increase in nacelle bay pressure.

BACKGROUND OF THE INVENTION

The present invention relates to nacelle designs for aircraft gasturbine engines and, more particularly, to a nacelle cooling andventilation system.

As a result of the high temperatures generated within the engine bay ofa nacelle housing a gas turbine engine, it is necessary to passsignificant amounts of cooling air through the bay to maintain thecomponents within acceptable temperatures and to meet industry standardfire safety criteria. Further, the lubricating oil for the propellergear box and generator requires a cooling system. Typically, cooling airis supplied through inlets in the nacelle housing into the engine bayand the stream of exhaust air from the engine induces the cooling airfrom the bay, and mixes it with the exhaust stream. Propeller gear boxand generator lubricating oils are cooled in air/oil heat exchangerswhich receive air from a scoop separate from the engine intake scoop#andrelease air overboard.

An example of a nacelle cooling system is disclosed in Hotz et al. U.S.Pat. No. 2,548,794. That patent discloses a nacelle in which the twingenerators of a twin engine nacelle are connected to receive cooling airfrom a duct which is separate from the engine inlet duct. An exhaustduct conveys the cooling air from the generators to the exhaust area ofthe nacelle which is aft of the engine bay so that the exhaust streamfrom the engines draws air through the generator.

In the Hotz et al. design, the engine bay itself is cooled by entry ofair through vents formed in the sides of the nacelle. Air is drawn inthrough those vents and is passed through the bay and into the inlet airto the engines. Disadvantage with such a design are: (1) any flammablefluid leakage in the engine bay might be carried into the engine inletwith a concurrent fire safety risk, and (2) air drawn in through theside vents decreases the overall efficiency of the engine in developingthrust and reduces the overall aerodynamic efficiency of the nacelle.

Accordingly, there is a need for a nacelle for a gas turbine engine inwhich both lubricating oil and the engine bay are cooled with externalair passed through the bay with minimum reductions in aerodynamicefficiency and in the amount of thrust developed by the engine.

SUMMARY OF THE INVENTION

The present invention is a nacelle cooling and ventilation system inwhich air enters the engine bay through a single nacelle opening and isused to cool oils in the generator and propeller gear box to ventilatethe engine bay. The air entering the engine bay is used for cooling andventilating the nacelle compartment, thereby eliminating the need foradditional louvers or vents in the nacelle during normal operation.Further, the air is exhausted through an engine exhaust eductor at theengine bay exit. The advantage of this system is that, by maintainingthe openings in the nacelle to a minimum, losses in thrust resultingfrom cooling the engine or oil systems is maintained at a minimum andall cooling air passes into the exhaust stream of the turbine enginewhere it contributes to thrust.

In a preferred embodiment, the system includes the start bleed flow inwhich pressurized air is ducted to an a ejector which supplements theengine exhaust eductor at the engine bay exit, rather than being dumpedoverboard as in prior art systems. This auxiliary ejector helps provideadequate air flow through the nacelle at low aircraft speeds such as atground runs or taxiing. Also in the preferred embodiment, the nacelleincludes two engine bay vent doors with actuators that open them in theevent that the oil temperature exceeds a predetermined limit, therebyincreasing airflow through the oil coolers and the engine bay. Further,the actuator is connected such that the doors are open to allow naturalventilation when the engine is shut down on the ground. The doors areconfigured to disconnect the actuator and pivot freely at apredetermined pressure differential across the doors so that excessiveengine bay pressure, as from an open bleed air duct, can escape throughthese vent openings thereby eliminating the need of separate blow-outdoors.

In another aspect of the preferred embodiment, the propeller gear boxand generator oil heat exchangers are combined in a single unit whichreceives cooling air from a single opening. A thermostaticallycontrolled set of louvers is mounted on the heat exchangers and thelouvers are oriented to direct air forwardly into the nacelle so thatall components of the turbine engine and propeller gear box have coolingair flowing over them before the air is drawn rearwardly to the exhausteductor. The thermostat incorporated into the air/oil heat exchangeropens the louvers to their fullest extent at a predetermined maximumtemperature, and closes the louvers to a minimum when oil temperaturefalls below a predetermined minimum. If an engine bay fire occurs, anaircraft control solenoid disconnects the control thermostat at engineshut down and the louvers shut completely to minimize air flow and toreduce the amount of fire extinguishing agent required.

Accordingly, it is an object of the present invention to provide acooling and ventilation system for a gas turbine nacelle which minimizesthe number of air inlet openings and air exhaust opening so that allexhausting air is dumped into the engine exhaust to promote high thrustefficiency; a cooling and ventilation system in which start bleed air isutilized to draw cooling air through the nacelle at low aircraft speedconditions; a cooling and ventilation system in, which an air/oil heatexchanger includes louvers that distribute air throughout the engine bayof the nacelle or shut air flow down in low temperature or fireconditions; a cooling and ventilation system which includes auxiliaryvent doors which open only under conditions of high propeller gear boxoil temperature or of engine shut-down on ground situations and act aspressure relief doors; and a cooling and ventilation system which isreliable, safe, which simplifies nacelle construction, is easy tomaintain and is relatively inexpensive to implement.

Other objects and advantages of the present invention will be apparentfrom the following description, the accompanying drawings and theappended claims.

BRIEF DESCRIPTION OF THE DRAWING

FIG. 1 is a somewhat schematic side elevation of a nacelle incorporatingthe preferred embodiment of the present invention, partially broken awayto reveal the engine bay;

FIG. 2 is a schematic side elevation of the gas turbine engine of FIG.1, taken from the opposite side of the engine bay;

FIG. 3 is a detail showing a top plan view of the gas turbine engine ofFIG. 1 in which the auxiliary exhaust eductor is partially broken away;

FIG. 4 is a detail schematic showing the air/oil heat exchanger of thenacelle of FIG. 1;

FIG. 5 is a detail showing the louver operation linkage of the heatexchanger of FIG. 4;

FIG. 6 is an electrical schematic detail of the circuit energizing thefire control solenoid of the present invention;

FIG. 7 is a detail showing the position of the bay vent doors of thenacelle of FIG. 1;

FIG. 8 is a schematic detail of the nacelle of FIG. 1 showing themechanism operating one of the bay vent doors shown in FIG. 7;

FIG. 9 is a cross section of the bell crank taken at line 9--9 of FIG.8; and

FIG. 10 is a detail cross section taken at line 10--10 of FIG. 9.

FIG. 11 is a sectional view taken along line 11--11 of FIG. 1illustrating the relative positions of the engine air inlet duct and thechin scoop and heat exchanger.

DETAILED DESCRIPTION

As shown in FIG. 1, the nacelle of the present invention, generallydesignated 10, houses a gas turbine engine 12 which drives a propellerassembly 14 through a propeller gear box 16. An engine air scoop 18 ispositioned above the propeller gear box 16 and conveys air forcombustion through S-shaped duct 19 to the engine 12, as generallydepicted by the hollow flow arrows in FIG. 1. A radially inner andaxially aft portion of S-shaped duct 19, which abuts the inlet to engine12, circumscribes an aft portion of fairing 21 which in turncircumscribes a shaft (not shown) which drivingly connects engine 12 topropeller gear box 16. A chin scoop 20 is positioned in a lower portionof the nacelle 10 and includes a duct 22 which conveys air to the enginebay 24 of the, as generally depicted by the solid flow arrows in FIG. 1.The engine bay 24 is designated fire zone A. The air entering bay 24through chin scoop 20 is directed forward to the area of propeller gearbox 16 before being drawn rearwardly, as best seen in FIG. 1 and asdiscussed in more detail subsequently, and additionally, as best seen inFIG. 11, flows radially outward around S-shaped duct 19 which preventsthis air from mixing with the engine inlet air used for combustion.

An air/oil heat exchanger assembly, generally designated 26 (also calledan oil cooler), is positioned within the duct 22 and is in fluidcommunication with the lubricating oil of the propeller gear box 16 andgenerator 28. The engine 12 includes an exhaust shroud 30 which enclosesthe exhaust conduit 31 of the engine (see FIG. 3) and forms an annularpassage 32 which allows air to flow from the bay 24 to the exhauststream Entrance 33 of annular passage 32 is adjacent to and in directfluid flow communication with bay 24.

As shown in FIG. 3, the cooling and ventilation system of the nacelle 10includes a start-bleed air exhaust assembly, generally designated 34.The start-bleed exhaust assembly 34 includes a conventional controlvalve 36 and actuator 38 and is connected to a midstage compressorportion of the engine 12 to adjust airflow from the axial compressor tomatch the centrifugal compressor at low engine speeds. A conduit 40extends from the valve 36 and is connected to a transition piece 42which is connected to an annular plenum 44. Annular plenum 44 isattached to and extends about the periphery of the exhaust shroud 30 andincludes an annular opening 46 covered by an annular flange 47. Theopening 46 which connects the plenum 44 to the exhaust shroud 30, whichterminates within the flange 47 so that air flowing into the plenumflows through the annular opening 46 to the space between the shroud andthe exhaust duct 31 and from there to the exhaust stream of the engine12.

As shown in FIGS. 4 and 5, air/oil heat exchanger assembly 26 includes apropeller gear box oil heat exchanger 50, a generator oil heat exchanger52 and a louver assembly 54. The propeller gear box and generator heatexchangers 50, 52 provide cooling of the lubricating oils utilized inthe propeller gear box 16 and generator 28. The heat exchangers 50, 52preferably each have a folded crossflow configuration and are aluminumplate-fin, air-to-oil units fabricated by brazing and welding. The corematrices of each heat exchanger 50, 52 are a brazed assembly consistingof alternate layers of oil and air fins separated by tubeplates. Doubleclosure bars are integrally brazed to the tubeplates to form the oil andair passages. These double bars on all the passages provide leakagecontainment for the fluids, avoiding air contamination by the oil.

The oil and air fins (not shown) in the units 50, 52 are made of thingauge aluminum alloy sheet. The air inlet side of the cooler 26incorporates a thicker gage fin at the inlet of each cooling air passageto protect the core matrices against foreign objects. As shown in FIG.4, the propeller gear box heat exchanger 50 includes inlet and outletlines 56, 58, respectively, connected to the propeller gear boxlubrication system (not shown) and the generator heat exchanger 52includes inlet and outlet lines 60, 62, respectively, connected to thegenerator lubrication system (not shown).

The louver assembly 54 includes a housing 64 which supports a pluralityof louvers 66 for pivotal rotation. Referring to FIG. 5, each louver 66is connected at its ends by a bearing mount 68 and the entire assemblyis positionable in unison by linkage 69 of conventional design which isactuated by displacing an actuator rod 70. The actuator rod 70 ispivotally connected to a link arm 72 which is fixed to an actuator arm74. The actuator arm 74 is loaded in a clockwise direction in FIG. 5 byreturn spring 76 and is displaced in a counter-clockwise direction bytemperature controlled actuator 78.

A fire shut-off solenoid 80 actuates a rod 81 which pivots a connectingpin 82 that extends between the temperature controlled actuator 78 andactuator arm 74. The solenoid 80 is energized by the aircraft powersupply and, as shown in FIG. 6, is controlled by propeller gear box lowpressure switch 83 (located on line 58) in nacelle 26 and engine firehandle switch 84. Switches 83, 84 are in series and both must be closedto energize solenoid 80.

In the preferred embodiment, the temperature controlled actuator 78includes a thermostatic organic wax element which expands and contractsto displace the pivot pin 82 sidewardly against the actuator arm 74.Counter-clockwise pivotal motion of the arm 74 causes the rod 70 to bedepressed upwardly in FIG. 5 to pivot the louvers to a generally openposition as shown in FIG. 4. Conversely, a decrease in temperaturecauses a sideward movement of the pin 82 allowing the arm 74 to pivotclockwise, thereby moving the rod 70 downwardly in FIG. 5 to close thelouvers 66. The actuator 78 is adjusted to position the louvers so thatoil temperature in the propeller gear box 16 is maintained below 165° F.and typically between 145° F. and 165° F.

As shown in FIGS. 5 and 6, in the event of a fire emergency, uponclosure of propeller gear box oil low pressure switch 83 indicating thatthe engine has been shut down by the flight crew, and actuation by theflight crew of the engine fire handle 84, the fire shut-off solenoid 80is energized to draw the rod 82 upwardly in FIG. 5 to pivot pin 84 outof contact with the actuator arm 74. This allows the return spring 76 todraw the arm 74 fully clockwise so that the louvers 66 assume a fullyclosed configuration. However, in normal operation the louvers 66 arepositioned by actuator 78 between a minimum slightly open and at amaximum are open at a maximum of the angle shown in FIG. 4. As shown inFIG. 1, this angular orientation directs air flowing through the chinscoop 20 forwardly within the engine bay 24 to ventilate the part of thenacelle forward of the oil cooler 26, as generally depicted by the solidflow arrows. The minimum position is consistent with fire safetycriteria of low residency time of flow over engine hot cases.

Further, the oil cooler 26 is positioned at an angle within the duct 22to prevent packing of the cooler by ice or snow when the associatedaircraft is flying in inclement weather. A drain opening 85 ispositioned rearwardly of the duct 22 and provides an exit forcondensation, other moisture collecting in the duct or dripping from thecooler 26. An oil drain port and tube 86 is connected to discharge oilleakage from the heat exchangers 50, 52 overboard.

As shown in FIG. 7, the cooling and ventilation system includes a pairof bay vent doors 87, 88 which are positioned on the upper surface ofthe nacelle 10 and forwardly of the infra red suppressor scoop 90. Thedoors 87, 88 are hinged in a longitudinal direction so that they openlaterally.

FIGS. 8, 9 and 10 show the details of the actuation mechanism, generallydesignated 92, of door 88, it being understood that a similar mechanismis provided for door 87. Actuation mechanism 92 includes an actuator inthe form of a cylinder motor 94 pivotally mounted to a bracket 96attached to the nacelle housing 97. The cylinder motor 94 includes a rod98 which is pivotally connected to a bell crank 100 and includes apiston 102 within a cylinder 104. The piston 102 is spring biased by anextension spring 106 to travel to the left in FIG. 8, thereby causingthe bell crank 100 to pivot clockwise about pivot connection 108attached to nacelle housing 97.

A chamber 110 in the cylinder 102 is selectively pressurized by air fromthe engine bleed air system 103 (see FIG. 2) which is supplied throughan actuator valve, generally designated 112.

Valve 112 includes a valve body 114 defining an inlet chamber 116 whichcommunicates with a valve chamber 118 by a conduit 120. Inlet chamber116 is connected to pressure line 122 which supplies pressurized airfrom the aircraft bleed air duct 123 of bleed air system 103 downstreamof shut off value 124 and actuator 125 (see FIG. 2). Conduit 126 in FIG.8 is connected to inlet chamber 116 and high pressure duct 127 of thebleed air system ducting of the engine 12 in FIG. 2 upstream of highpressure valve 128 and actuator 129.

A spherical shut-off element 130 is captured within the inlet chamber116. When pressurized air from high pressure component 128, conveyedthrough conduit 126, enters inlet chamber 116 and is higher in pressurethan pressurized air entering through conduit 122 from environmentalcontrol system conduit 124, the element 130 is displaced to shut offconduit 122 so that compressed air is directed through passageway 132 tovalve chamber 118. Conversely, when the bleed air system 103 is shutdown, the pressure through conduit 126 drops and air pressure throughconduit 122 from the aircraft bleed system exceeds that entering throughconduit 126, the element 130 is displaced to the right in FIG. 8 so thatcompressed air is directed through passageway 132. Consequently, valve112 maintains the actuator 92 pressurized and the doors 87, 88 closedwhen the engine 12 is shut down and the aircraft bleed system ispressurized, thereby reducing installation drag during an in flight shutdown condition.

Valve chamber 118 includes a pintle 134 which is biased by spring 136 tothe closed position shown in FIG. 8. In this position, air in chamber110 travels through conduit 132 to be dumped through exit port 138,thereby allowing the spring 106 to displace the cylinder 102 in rod 98to the left in FIG. 8 to open the door 88.

A control system 140 actuates solenoid 142 to displace the pintle 134 tothe right, thereby closing the exit opening 138 and allowing compressedair entering inlet chamber 116 to flow through passageway 120 andconduit 120 into chamber 110, thereby displacing the piston 102 and rod98 to the right in FIG. 8 to maintain the door 88 in the closedconfiguration shown. Control system 140 includes normally closed switch144 and normally open switch 146. Normally closed switch 144 is an oiltemperature switch on outlet line 58 of heat exchanger 50 (see also FIG.4). Switch 144 is set to open when propeller gear box oil temperatureexceeds 170° F. In this condition, both switches 144, 146 are open, thesolenoid 142 is not energized and the vent door 88 opens since the airin chamber 110 is dumped through exit opening 138 in valve 112.

Switch 146 is normally open and is a pressure switch mounted on duct 40which closes when the start bleed duct is pressurized. Consequently, thevent doors 87, 88 are not opened at engine idle conditions so thatshroud 30 draws air only from chin scoop 20 (see FIG. 1).

The bell crank 100 includes a pair of brackets 148, 150 orientedparallel to each other and connected by cross webbing 152. The brackets148, 150 are spaced apart sufficiently to form a recess 154 sized toreceive the flange 156 connected to the door 88. The flange 156 isconnected to pivot 108. The flange 156 includes detent slots 158 on bothsides which receive the pins 159 of detents 160 carried in brackets 148,150. Detent springs 162 are adjusted by screws nut 164 and lock nuts 165such that the retentive force exerted by the detents 160 are sufficientto maintain the door 88 in a closed position when the pressuredifferential across the door 88 is below a predetermined valuecompatible with the structural integrity of the nacelle cowlings.

When the pressure within the bay 24 exceeds this value, the pressure onthe door 88 causes the flange 156 to disengage from the detents 160,permitting the door 88 to pivot about pivot 108 to an openconfiguration. The door 88 is prevented from opening beyond a maximumdesired angle by a lanyard 166 which is attached to a door bracket 168and a bracket 170 on the nacelle housing 97. The flange 156 will notautomatically re-engage the crank 100 but will remain partiallydisengaged to signal to ground crews that a blow out event has occurredin flight.

The operation of the nacelle cooling and ventilating system is asfollows. When the engine 12 is at low speeds, or during starting, thestart bleed valve 36 is open to pressurize the start bleed duct 40.Switch 146 is closed which actuates valve 112 to close the vent baydoors 87, 88. At this time, the louvers 66 of the oil cooler 26 are inthe minimum open position, but some air does flow into bay 24 as aresult of the drawing power of the start bleed exhaust assembly 34. Asair speed increases and oil temperatures increase, valve 78 opens thelouvers 66 such that air entering the chin scoop 20 and travelingthrough duct 22 is directed into the bay 24 and forward to the propellergear box area 16, before being drawn rearwardly through the annularopening 32 of the exhaust shroud 30.

The temperature controlled actuator 78 continually positions the louvers66 to provide sufficient air flow to cool the oil in the heat exchanger26 for the propeller gear box 16 and generator 28, and modulating airflow through the bay 24 for cooling and ventilating the engine bay 24 ofthe engine 12. During take off and cruising speeds, the start bleedexhaust assembly is inactive; however, air flow through the chin scoop20 is sufficient to provide for the cooling needs of the oil lubricatingthe propeller gear box 16 and generator 28 due to the ram air effect onchin scoop 20 so that the additional drawing force of the start bleedejector 34 is not required. All cooling air, during both start up andcruising or operational conditions, is vented into the tailpipe exhaustto add to the thrust produced by the engine 12.

At above idle engine speeds, the start bleed valve 36 is closed so thatswitch 146 is open. Whenever the propeller gear box oil temperature isbelow its maximum limit, switch 144 is closed which directs pressurizedair from the bleed system to pressurize cylinder motor 94 to close thevent bay doors 87, 88. However, should a propeller gear boxovertemperature condition exist, switch 144 will open which willde-energize the solenoid 142 to block air flow from the bleed system 103and, at the same time, allow the air within the cylinder motor 94 to bedumped through the exit opening 138, thereby allowing spring 106 to openthe bay vent doors 87, 88. If an overpressure condition occurs withinthe bay 24, the vent doors 87, 88 open in response to the pressuredifferential across them and allow for a release of pressure.

While the form of apparatus herein described constitutes a preferredembodiment of this invention, it is to be understood that the inventionis not limited to this precise form of apparatus, and that changes maybe made therein without departing from the scope of the invention.

What is claimed is:
 1. In a nacelle having a turbine engine in a baythereof, an engine inlet for supplying combustion air to said engine, anexhaust shroud forming an annular passage with an exhaust conduit ofsaid engine for evacuating air from within said engine bay, a chin scoopfor receiving cooling air and an air/oil heat exchanger receivingcooling air from said chin scoop, a cooling and ventilation systemcomprising:duct means for discharging air from said heat exchanger intosaid bay wherein said discharged air cools said engine and is drawnthrough said annular passage into said engine exhaust, whereby thrustrecovery is minimized and overboard dumping of cooling air is minimized;wherein said annular passage includes an entrance adjacent to and indirect fluid flow communication with said bay; wherein said bay issubstantially enclosed by said nacelle wherein effectively all coolingair entering said bay enters from said chin scoop, wherein the coolingair provided for use in said air/oil heat exchanger and for cooling saidengine is provided with minimum reductions in aerodynamic efficiency;said cooling and ventilation system further comprising means forregulating air flow from said heat exchanger through said bay such thatnacelle cooling and ventilation flows are held to a minimum; whereinsaid regulating means includes adjustable louver means mounted on saidheat exchanger and temperature sensing means for positioning said louvermeans to increase or decrease air flow to said nacelle whereby baytemperature of oil cooled by said heat exchanger is maintained within apredetermined range; and wherein said louver means is oriented todischarge cooling air toward a forward portion of said nacelle beforethe cooling air is drawn rearwardly through said annular passage.
 2. Thesystem of claim 1 wherein said temperature sensing means includes springreturn means for maintaining said louver means in a substantially closedconfiguration in both a power failed condition and an engine shut-downcondition.